1. Field of the Invention
The present invention relates to compressors and, more particularly, to a multi-stage compressor rotor for a gas turbine engine.
2. Description of the Prior Art
Multi-stage compressors having an axial-flow stage followed by a centrifugal stage are known in the art. Such multi-stage compressors typically comprise an axial-flow rotor and a centrifugal rotor or impeller having respective disc-like portions connected to each other by means of bolts or the like. The axial-flow rotor and the centrifugal rotor are formed separately and then connected to each other with an axial gap between respective arrays of circumferentially spaced-apart blades thereof. The forging required to form the axial-flow rotor and the centrifugal rotor is considerable and the axial gap between their respective arrays of blades might result in unsynchronized deflection as the air passes from one stage to the next and, thus, adversely affect the overall aerodynamic performance of the multi-stage compressor.
Therefore, there is a need for a new multi-stage compressor rotor requiring less forging while having improved aerodynamic performances.
It is therefore an aim of the present invention to provide a new multi-stage compressor rotor having improved aerodynamic performance.
It is also an aim of the present invention to improve the growth potential of a compressor rotor.
It is a further aim of the present invention to provide a multi-stage compressor rotor of relatively light weight construction.
It is a still further aim of the present invention to provide a multi-stage compressor which is relatively simple and economical to manufacture.
Therefore, in accordance with the present invention, there is provided a multi-stage compressor rotor for a gas turbine engine, comprising an axial-flow rotor followed by a centrifugal rotor, said axial-flow rotor and said centrifugal rotor being bonded together to form a unitary dual flow impeller having blades with united axial-flow and centrifugal stage sections.
In accordance with a further general aspect of the present invention, there is provided a multi-stage compressor rotor for a gas turbine engine, comprising an axial-flow rotor followed by a centrifugal rotor, said axial-flow rotor and said centrifugal rotor being provided with respective arrays of circumferentially spaced-apart blades, wherein each blade of said centrifugal rotor extends in continuity from a corresponding blade of said axial-flow rotor to a discharge edge thereof.
In accordance with another general aspect of the present invention, there is provided a dual flow impeller for a gas turbine engine, comprising a disc-like member having front and rear sections bonded together, an array of circumferentially spaced-apart blades defined in said front and rear sections, each said blade having a continuous blade profile including an axial-flow inducing stage section followed by a centrifugal-flow stage section.